This invention relates to the cooling of aerofoils in a gas turbine engine.
The performance of the simple gas turbine engine cycle, whether measured in terms of efficiency or specific output is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbine at the highest possible temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust (eg engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an uncooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of the blades' materials used and in some engine designs the intermediate pressure (IP) and low pressure (LP) turbines are also cooled. During its passage through the turbine the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the HP stage(s) through the IP and LP stages towards the exit nozzle.
Internal convection and external films are the prime methods of cooling the aerofoils. HP turbine nozzle guide vanes (NGVs) consume the greatest amount of cooling air on high temperature engines. HP blades typically use about half of the NGV flow. The IP and LP stages downstream of the HP turbine use progressively less cooling air.
Blades and vanes are cooled by using high pressure (HP) air from the compressor that has by-passed the combustor and therefore is relatively cool compared to the temperature of the working gas in the gas path. Typical cooling air temperatures are between 700 and 900 K. Gas path temperatures can be in excess of 2100 K. The cooling air extracted from the compressor and used to cool hot turbine components is not used fully to extract work from the turbine. Extracting coolant flow therefore has an adverse effect on the engine operating efficiency and it is thus important to use this cooling air as effectively as possible.
Historically, in a simple internal air cooling system the cooling air has been passed through an aerofoil in the radial direction from blade root to tip, often with provision for effusion cooling of the leading and trailing edges. The efficiency of such systems is limited because high cooling efficiency is obtained with a passage having a high length to diameter ratio. Other factors have to be considered, the minimum passage “diameter” is limited, for example by the manufacturing process utilised and by aerofoil weight, and the length is dictated by the size of the component. More recently improved casting technology has enabled the use of a multi-pass cooling arrangement, where the flow is passed up and down the component a number of times and has resulted in higher efficiencies than those obtained by a simple radial system.
One characteristic, although not the sole one, affecting the performance of an internal cooling system is the amount of heat absorbed by the coolant. The rate at which the coolant absorbs heat is dependent on the temperature difference between the surrounding metal surface and the coolant. Consequently systems tend to function less well towards the end of a multi-pass arrangement, and the temperature of the metal surrounding the cooling passage changes from relatively cold near the inlet to relatively hot near the exit. The resulting temperature gradients are undesirable but to some extent unavoidable.
Prior art attempts to mitigate the effects of the above-mentioned characteristic feature of internal air cooling systems have employed a plurality of multi-pass cooling arrangements. Examples of such arrangements are described in earlier published patents numbers in order of publication date: GB 1,188,401 of 1970; U.S. Pat. No. 4,818,178 of 1987; JP8246802 and JP 8260901 both of 1996; GB 2,322,167 of 1998; EP 1,327,747 of 2003 and EP 1,319,803 of 2004. In all of these documented prior arrangements the internal air-cooling systems illustrated comprise multi-pass cooling arrangements arranged in tandem, that is in the chordal direction of the blade one of the cooling arrangements is positioned towards the leading edge and a second behind it towards the trailing edge. JP 60198305 published in 1985 also shows an internal air-cooling system employing a plurality of multi-pass cooling arrangements comprising a tandem pair adjacent the pressure surface of the blade and another tandem pair adjacent the suction surface of the blade. It is common ground to all of these arrangements that the inlet ends of the multi-pass serpentine passages to which cooling air is supplied are located adjacent one another and towards a point midway between the leading and trailing edges. Thus, the cooling arrangements absorb heat most efficiently towards the centre of the blade where the cooling air has most heat capacity and operate less efficiently adjacent the leading and trailing edges where some of the heat capacity of the air has already been taken up. As a result the cooling arrangements tend to exaggerate a temperature gradient within the metal of the blade by preferentially cooling the centre of the blade while at the same time allowing the temperature of the leading and trailing edges to increase.